Direct Numerical Simulation of Film Cooling in Hypersonic Boundary-Layer Flow

Abstract

Effusion cooling by discrete slits and holes in various laminar zeropressure gradient super-and hypersonic boundary layers is investigated using direct numerical simulation. For an adiabatic Mach-6 boundary layer it was found that slits are better than holes due to the lower blowing velocity. Slit blowing causes a destabilisation of 2 nd -mode disturbances, and a complete stabilisation of 1 st modes despite the generated maxima of the spanwise vorticity inside the boundary layer. Hole blowing gives rise to counter-rotating streamwise vortices, with a noticeable laminar-flow destabilisation only for large spanwise hole spacings. Computational aspects such as the performance of the numerical code on the supercomputer NEC SX-8 and optimization of the code are also discussed.

Effects of Wall-Temperature Conditions on Effusion Cooling in a Supersonic Boundary Layer

“…It was found that aligned holes and a large spanwise spacing jeopardize laminarity of the flow. Nevertheless, no transition has been observed in these investigations, due to the weaker growth of the disturbance waves in supersonic flows as well as a weaker wall shear, compared with the incompressible case [20]. For this reason, staggered holes and a small spanwise spacing are preferable, which is also good for effusion cooling, because of the lower blowing velocities and a more homogeneous spanwise distribution.…”

“…The flow physics of effusion-hole configurations, aligned or staggered rows of holes, has also been analyzed for an adiabatic laminar Mach 6 boundary layer by Linn and Kloker [17,20]. It was found that aligned holes and a large spanwise spacing jeopardize laminarity of the flow.…”